BACKGROUND INFORMATION
1. Field:
[0001] This disclosure generally relates to composite structures, especially fiber reinforced
resin laminates, and deals more particularly with a hybrid composite having a composite-to-metal
joint, as well as to a bonded metal laminate used in the joint.
2. Background:
[0002] A composite structure having composite-to-metal joints is known from
US 2012/0045606 A1.
[0003] Bonding techniques are often used to assemble composite structures. In applications
where the composite structure also requires fasteners, the local thickness or gauge
of the structure surrounding the fastener may need to be increased in order to withstand
loads transmitted through the fastener joint. As the local thickness of the structure
increases, the fastener may need to be lengthened, thereby adding weight to the structure.
Additionally, the increased local thickness of the structure may increase the eccentricity
of the load path across the fastener joint, which may place undesired bending loads
on the fastener.
[0004] One solution to the problems mentioned above consists of attaching metal fittings
to the composite structure in the area of the fasteners. These metal fittings may
be formed of titanium or similar metals that may not substantially chemically react
with carbon fiber reinforced composites in which they are in contact. Titanium fittings,
however may be relatively expensive, particularly when it is necessary to form them
into complex shapes.
[0005] Accordingly, there is a need for a composite resin-to-metal joint that may be used
to connect substantially all metal fittings with substantially all composite resin
structures, which is relatively inexpensive and easy to manufacture, and which may
withstand loads transferred around fastener connection points. There is also a need
for a composite resin-to-metal joint that substantially avoids chemical reactions
between the all metal fitting and the all composite resin structure. Also, there is
a need for a composite-to-metal joint that may reduce residual stresses in the joint
following a thermal curing. Further there is a need for a bonded metal laminate that
may be used in the joints and in other applications where additional strength and
durability are required.
SUMMARY
[0006] The above-mentioned need is overcome by a structure in accordance with claim 1, which
defines the invention.
[0007] The present disclosure provides a hybrid-type composite structure that includes a
fiber reinforced resin composite-to-metal joint that may be used to connect a substantially
all-metal fitting with a substantially all composite resin structure or a different
structure. The joint provides a transition between the composite and metallic structures
that is suitable for use in higher performance applications, such as aerospace vehicles.
This transition from a substantially all composite to a substantially all metal material
may reduce or eliminate the possibility of corrosion and/or problems stemming from
eccentricity. During lay-up of the composite structure, relatively thin, flexible
metal sheets of metal are substituted for a number of composite plies, and the transition
from composite plies to metal sheets occurs at staggered locations so as to provide
adequate load transfer from the composite portion to the metal portion. The staggered
transition results in an interleaving between the composite plies and the metal sheets
and creates multiple bond lines that may reduce the occurrence and/or propagation
of cracks or disbonds in the joint. An adhesive placed between the metal sheets binds
and unitizes the sheets into a nearly solid metal fitting.
[0008] The composite-to-metal joint may be configured as a finger type, step lap joint in
order to reduce residual stresses that may be induced in the joint during cooling
of the hybrid composite structure following a thermal cure cycle. The bonded metal
sheets employed in the joint form a metal laminate that may be used in a variety of
other applications, and which exhibits improved performance compared to monolithic
metal structures. In some applications, the composite-to-metal joint utilizing the
metal laminate may be used to reinforce an edge of a composite structure or to reinforce
an area of a composite structure around fasteners. Additional advantages of the disclosed
composite-to metal joint may include improved joint robustness, reduced weight, improved
safety, less maintenance, weight savings, improved inspectability, strength improvements,
and reduced manufacturing costs. The disclosed metal laminate used in the composite-to-metal
joint may enable a structure to have weight and fatigue characteristics of composite
resin laminates while providing the strength and durability of a metal structure.
The composite-to-metal joint may reduce or avoid the need for machined end-fittings
for some composite resin structure applications. A shorter bond length resulting from
use of the disclosed joint may minimizes residual (or cured in) stresses due to CTE
(coefficient of thermal expansion) mismatch between the metallic and composite materials
forming the joint, and may also benefit the in-service performance of the joint where
service temperatures can vary 225 degrees F or more.
[0009] According to the present disclosure, a metal structure is provided that exhibits
improved strain performance. The metal structure comprises at least a first metal
laminate including a first plurality of metal sheets bonded together. The metal structure
further comprises a plurality of layers of a bonding adhesive forming adhesive bonds
between the metal sheets. The metal laminate includes at least one through hole therein
adapted to receive a fastener. The metal structure may further comprise a second metal
laminate including a second plurality of metal sheets bonded together, and at least
one fastener joining the first and second metal laminates together.
[0010] According to the present disclosure, an integrated attachment fitting is provided
for a structure. The attachment fitting comprises a composite resin portion, a metal
portion, and a composite-to-metal joint between the composite resin portion and the
metal portion. The composite resin portion includes a plurality of fiber reinforced
resin plies, and the metal portion includes a plurality of metal sheets bonded together.
The composite-to-metal joint includes overlapping steps between the fiber reinforced
resin plies and the metal sheets. The composite-to-metal joint may comprise a finger
joint. In one application, the structure may comprise an aircraft vertical stabilizer,
and the metal portion may be a metal laminate attachment lug having a through-hole
therein adapted to receive a bolt for attaching the lug to an aircraft fuselage. The
composite resin portion forms part of the aircraft vertical stabilizer. In another
application, the structure may be an aircraft wing, and the metal portion is a metal
laminate having a plurality of through-holes therein adapted to receive fasteners
for attaching the wing to a center wing box on an aircraft fuselage. The composite-to-metal
joint may be one of a finger lap joint, a tapered lap joint, a vertical lap joint,
and a lap joint having a variable overlap. In a further application, the structure
may be a rotor blade having a root adapted to be attached to a rotating hub, and the
metal portion includes a metal laminate located at the root, wherein the metal laminate
has a through-hole therein adapted to receive a retention bolt for retaining the rotor
blade on the rotating hub. In still another application, the composite-to-metal joint
is an overlapping splice joint adapted to join two fuselage sections of an aircraft.
[0011] According to the present disclosure, a fastener reinforcement is provided for reinforcing
an area of a multi-ply composite structure. The fastener reinforcement comprises a
metal laminate including a plurality of metal sheets bonded together, wherein the
metal laminate has a through-hole adapted to receive a fastener therein. The fastener
reinforcement further comprises a composite-to-metal joint between the metal laminate
and the composite structure. The meal sheets have edges that are interleafed with
the plies of the composite structure.
[0012] According to the present disclosure, a method is provided of fabricating a composite
structure, comprising assembling at least a first stack of metal sheets, and laminating
the first stack of metal sheets together by placing a layer of adhesive between each
of the metal sheets. The method further comprises assembling a second stack of metal
sheets, laminating the second stack of metal sheets together by placing a layer of
adhesive between each of the metal sheets, and fastening the first and second stacks
of metal sheets by passing fasteners through the first and second stacks of metal
sheets.
[0013] According to the present disclosure, a method is provided of reinforcing an area
of a composite laminate containing a fastener passing through the thickness of the
composite laminate. The method comprises integrating a multi-ply metal laminate into
the area of the composite laminate to be reinforced, and forming a through-hole in
the metal laminate for receiving the fastener. Integrating the metal laminate is performed
by interleafing plies of the metal laminate with plies of the composite laminate to
form a finger joint between the metal laminate and the composite laminate. According
to the present disclosure, a method is provided of reinforcing an edge of a multi-ply
fiber reinforced resin laminate. The method comprises joining a metal laminate to
the resin laminate along the edge of the resin laminate. Joining the metal laminate
to the resin laminate is performed by interleafing edges of the plies of the metal
laminate and the resin laminate. The interleafing may be performed in a manner to
form a finger joint between the metal laminate and the resin laminate.
[0014] According to an aspect of the present disclosure there is provided a metal structure
exhibiting improved strain performance, comprising at least a first metal laminate
including a first plurality of metal sheets bonded together. Advantageously the metal
structure further comprises a plurality of layers of a bonding adhesive forming adhesive
bonds between the metal sheets. Advantageously the first metal laminate includes a
least one through hole therein adapted to receive a fastener therein. Advantageously
the metal structure further comprises a second metal laminate including a second plurality
of metal sheets bonded together and at least one fastener joining the first and second
metal laminates together.
[0015] According to an aspect of the present disclosure there is provided an integrated
attachment fitting for a structure, comprising a composite resin portion, a metal
portion, and a composite-to-metal joint between the composite resin portion and the
metal portion. Advantageously the composite resin portion includes a plurality of
fiber reinforced resin plies, and the metal portion includes a plurality of metal
sheets bonded together. Preferably the composite-to-metal joint includes overlapping
steps between the fiber reinforced resin plies and the metal sheets. Preferably the
composite-to-metal joint is a finger joint. Advantageously the structure is an aircraft
vertical stabilizer and the metal portion is a metal laminate attachment lug having
a through hole therein adapted to receive a bolt for attaching the lug to an aircraft
fuselage. Preferably the composite resin portion forms part of the aircraft vertical
stabilizer. Advantageously the structure is an aircraft wing and the metal portion
is a metal laminate having a plurality of through holes therein adapted to receive
fasteners for attaching the wing to a center wing box on an aircraft fuselage. Advantageously
the composite-to-metal joint is one of a finger lap joint, a tapered lap joint, a
vertical lap joint, and a lap joint having a variable overlap. Advantageously the
structure is a rotor blade having a root adapted to be attached to a rotating hub,
and the metal portion includes a metal laminate located at the root, the metal laminate
having a through hole therein adapted to receive a retention bolt for retaining the
rotor blade on the rotating hub. Preferably the composite portion is elongate and
extends outwardly from the root. Advantageously the structure is an aircraft wing
and the metal portion is a metal laminate having a plurality of through holes therein
adapted to receive fasteners for attaching the wing to a center wing box on an aircraft
fuselage. Advantageously the composite-to-metal joint is a lap joint adapted to join
two fuselage sections of an aircraft.
[0016] According to an aspect of the present disclosure there is provided a fastener reinforcement
for reinforcing an area of a multi-ply composite structure, comprising a metal laminate
including a plurality of metal sheets bonded together, the metal laminate having a
through hole therein adapted to receive a fastener therein and a composite-to-metal
joint between the metal laminate and the composite structure. Advantageously the metal
sheets have edges that are interleafed with edges of the plies of the composite structure.
[0017] According to an aspect of the present disclosure there is provided a method of fabricating
a composite structure, comprising assembling at least a first stack of metal sheets
and laminating the first stack of metal sheets together by placing a layer of adhesive
between each of the metal sheets. Advantageously the method further comprises assembling
a second stack of metal sheets, laminating the second stack of metal sheets together
by placing a layer of adhesive between each of the metal sheets and fastening the
first and second stacks of metal sheets together by passing fasteners through the
first and second stacks of the metal sheets.
[0018] According to an aspect of the present disclosure there is provided a method of reinforcing
an area of a composite laminate containing a fastener passing through the thickness
of the composite laminate, comprising integrating a multi-ply metal laminate into
the area of the composite laminate to be reinforced and forming a through hole in
the metal laminate adapted to receive the fastener therein. Advantageously, integrating
the metal laminate into the composite laminate is performed by interleafing plies
of the metal laminate with plies of the composite laminate to form a finger joint
between the metal laminate and the composite laminate.
[0019] According to an aspect of the present disclosure there is provided a method of reinforcing
an edge of a multi-ply fiber reinforced resin laminate, comprising joining a metal
laminate to the resin laminate along the edge of the resin laminate. Advantageously,
joining the metal laminate to the resin laminate is performed by interleafing edges
of the plies of the metal laminate and the resin laminate. Preferably the interleafing
is performed in a manner to form a finger joint between the metal laminate and the
resin laminate.
BRIEF DESCRIPTION OF THE ILLUSTRATIONS
[0020]
FIG. 1 is an illustration of a sectional view of a composite structure having a composite-to-metal
joint.
FIG. 2 is an illustration of a perspective view of the composite structure including
the composite-to-metal joint.
FIG. 3 is an illustration of a perspective view of the area designated as FIG. 3 in
FIG. 2.
FIG. 4 is an illustration of a cross sectional view of the joint, better showing interleaving
between composite plies and the metal sheets.
FIG. 5 is an illustration of a cross sectional view of two separated layers of the
joint shown in FIG. 4, also showing the application of a film adhesive on the metal
sheets.
FIG. 6 is an illustration of an enlarged, cross sectional view of a portion of the
joint formed by the two layers shown in FIG. 5.
FIG. 7 is an illustration of a broad flow diagram of a method of making a composite
structure having the composite joint shown in FIGS. 2-4.
FIG. 8 is an illustration of a flow diagram showing additional details of the method
shown in FIG. 7.
FIG. 9 is a flow diagram of another method of making a composite structure having
the composite joint shown in FIGS. 2-4.
FIG. 10 is an illustration of a perspective view of a composite-to-metal finger joint
having a relatively shallow double taper.
FIG. 11 is an illustration similar to FIG. 10 but showing a composite-to-metal finger
joint having a relatively steep taper.
FIG. 12 is an illustration of a sectional view of a composite-to-metal joint having
a single taper.
FIG. 13 is an illustration similar to FIG. 12 but illustrating a composite-to-metal
joint having a reversed single taper.
FIG. 14 is an illustration of a cross sectional view of a composite-to-metal finger
joint having a symmetric double taper.
FIG. 15 is an illustration similar to FIG. 14 but illustrating a symmetric reversed
double taper finger joint.
FIG. 16 is an illustration of a cross sectional view of a vertical composite-to-metal
finger joint, according to the invention.
FIG. 17 is an illustration of a cross sectional view of a composite-to-metal finger
joint having variable overlap between the plies.
FIG. 18 is an illustration of a plan view of a composite structure having a laminated
metal reinforcement around a fastener.
FIG. 19 is an illustration of a cross sectional view taken along the line 19-19 in
FIG. 18.
FIG. 20 is an illustration of an exploded, perspective view of a typical aircraft
employing composite-to-metal joints.
FIG. 21 is an illustration of a sectional view taken along the line 21-21 in FIG.
20, showing a typical composite-to-metal joint between fuselage sections.
FIG. 22 is an illustration of a perspective view of a composite-to-metal joint between
an aircraft wing and a center wing box.
FIG. 23 is an illustration of a perspective view of a portion of a skin of the wing
box shown in FIG. 22.
FIG. 24 is an illustration of the area designated as FIG. 24 in FIG. 22.
FIG. 25 is an illustration of a perspective view of an aircraft vertical stabilizer,
parts being broken away in section for clarity.
FIG. 26 is an illustration of a side view showing attachment of the stabilizer shown
in FIG. 25 to a fuselage using a lug containing a composite-to-metal joint.
FIG. 27 is an illustration of a side view of a forward portion of an aircraft, illustrating
a hatchway reinforced by a composite-to-metal joint.
FIG. 28 is an illustration of a sectional view taken along the line 28-28 in FIG.
27.
FIG. 29 is an illustration of a perspective view of a helicopter.
FIG. 30 is an illustration of a perspective view of a rotor assembly of the aircraft
shown in FIG. 29.
FIG. 31 is an illustration of the area designated as FIG. 31 in FIG. 30.
FIG. 32 is an illustration of a cross sectional view of a bonded metal laminate.
FIG. 33 is an illustration of a cross sectional view of two bonded metal laminates
joined together by fasteners.
FIG. 34 is a flow diagram showing a method fabricating the bonded metal laminate shown
in FIG. 32.
FIG. 35 is an illustration of a flow diagram of a method of reinforcing a composite
laminate containing a fastener.
FIG. 36 is an illustration of a flow diagram of aircraft production and service methodology.
FIG. 37 is an illustration of a block diagram of an aircraft.
DETAILED DESCRIPTION
[0021] Referring first to FIG. 1, a hybrid composite structure 20 includes a composite resin
portion 22 joined to a metal portion 24 by a transition section 25 that includes a
composite-to-metal joint 26. In the illustrated example, the composite structure 20
is a substantially flat composite sheet, however depending upon the application, the
structure 20 may have one or more curves, contours or other geometric features. For
example, composite structure 20 may comprise an inner and/or outer contoured skin
20 of an aircraft (not shown) which is secured to a frame portion 28 of the aircraft
by means of a lap joint 30 and fasteners 32 which pass through the composite structure
20 into the frame portion 28.
[0022] The frame portion 28 may comprise a composite, a metal or other rigid material, and
the metal portion 24 of the structure 20 may serve as a rigid metal fitting 24 that
is suited to transfer a range of loads and types of loadings between the frame portion
28 and the composite portion 20. As will be discussed below in more detail, the metal
portion 24 may comprise any of various metals such as, without limitation, titanium
that is substantially nonreactive to and compatible with the composite portion 22
and the frame portion 28. In one practical embodiment for example, and without limitation,
the composite resin portion 22 may comprise a carbon fiber reinforced epoxy, the metal
portion 24 may comprise a titanium alloy, and the frame 28 may comprise an aluminum
alloy or a composite. The transition section 25 and the joint 26 are strong enough
to carry the typical range and types of loads between the composite resin portion
22 and the metal portion 24, including but not limited to tension, bending, torsion
and shear loads. Although the illustrated transition section 25 and joint 26 are formed
between an all composite resin portion 22 and the all metal portion 24, it may be
possible to employ them to join two differing composite structures (not shown) or
two differing metal structures (not shown).
[0023] Referring to FIGS. 1-4, a layup of composite material plies 35 is terminated at a
interface location 39 referred to later herein as a transition point 39, where a metal
sheet or ply 37 of the substantially the same thickness as the composite material
plies 35 continues to the metal edge 24a of the metal portion 24, and the layup is
repeated with a composite-to-metal interface 39 that is staggered toward the metal
edge 24a from the prior interface location 39 and includes a ply of structural metal
adhesive 45 (see FIGS. 5 and 6) between the metal plies 37, with the next composite-to-metal
interface 39 staggered away from the metal edge 24a to produce a nested splice 27.
This staggered interface stacking, which produces nested tabs 29 (see FIG. 3), is
continued to the full thickness of the hybrid composite structure 20 with none of
the composite plies 35 extending fully to the metal edge 24a of the all metal portion
24
[0024] Referring now also to FIGS. 2-4, the composite portion 22 of the structure 20 comprises
a laminated stack 34 of fiber reinforced resin plies 35, and the metal portion 24
of the structure 20 comprises a stack 36 of metal sheets or plies 37 that are bonded
together to form a laminated, substantially unitized metal structure. As shown in
FIGS. 5 and 6, the composite plies 35 and the metal sheets 37 are arranged in layers
38. Each of the layers 38 comprises one or more of the composite plies 35 in substantially
edge-to-edge abutment with one of the metal sheets 37. Thus, each of the layers 38
transitions at a point 39 from a composite i.e. composite resin plies 35, to a metal,
i.e. metal sheet 37.
[0025] The transition points 39 are staggered relative to each other according to a predetermined
lay-up schedule such that the plies 35 and the metal sheets 37 overlap each other
in the transition section 25 (FIG. 1). Staggering of the transition points 39 creates
multiple bond lines that may reduce the occurrence and/or propagation of cracks or
disbonds in the joint 26. The staggering of the transition points 39 also results
in a form of interleaving of the composite plies 35 and the metal sheets 37 within
the joint 26 which forms a nested splice 27 between the all composite portion 22 and
the all metal portion 24. This nested splice 27 may also be referred to as a finger
bond 26, a finger joint 26 or a multiple step lap joint 26. The adjacent ones of the
transition points 39 are spaced from each other in the in-plane direction of the structure
20 so as to achieve a bonded joint 26 that exhibits optimum performance characteristics,
including strength and resistance to disbonds and propagation of inconsistencies such
as cracks. In the illustrated example, the nested splice 27 forming the joint 26 is
a form of a double finger joint 26 in which the transition points 39 are staggered
in opposite directions from a generally central point 55 of maximum overlap. However,
as will be discussed blow in more detail, other joint configurations are possible
including but not limited to a single finger joint in which the multiple transition
points 39 are staggered in a single direction.
[0026] The composite plies 35 may comprise a fiber reinforced resin, such as without limitation,
carbon fiber epoxy, which may be in the form of unidirectional prepreg tape or fabric.
Other fiber reinforcements are possible, including glass fibers, and the use of non-prepreg
materials may be possible. The composite plies 35 may have predetermined fiber orientations
and are laid up according to a predefined ply schedule to meet desired performance
specifications. As previously mentioned, the bonded sheets 37 may comprise a metal
such as titanium that is suitable for the intended application. In the illustrated
example, the stack 36 of metal sheets 37 has a total thickness
t1 which is generally substantially equal to the thickness
t2 of the laminated stack 34 of plies 35. In the illustrated example however,
t2 is slightly greater than
t1 by a factor of the thickness of several overwrap plies 43 on opposite sides of the
stack 37.
[0027] The use of a multiple step lap joint 26 may increase the bond area along the length
of the transition section 25, compared to a scarf type joint or other types of joints
which may require a longer length transition section 25 in order to achieve a comparable
bond area between the composite resin portion 22 and the metal portion 24. Following
thermal curing, cooling of the hybrid composite structure 20 may result in residual
stresses in the joint 26 due to a mismatch between the coefficient of thermal expansion
(CTE) of the composite resin portion 22 and the metal portion 24. The amount of thermal
expansion during curing is a function of the CTE of the composite resin portion 22
and the metal portion 24, as well as the length of the transition section 25. Use
of the step lap joint 26, rather than a scarf type or other type of joint may reduce
the amount of these residual stresses because of the reduction in the length of the
transition section 25 that is needed to obtain a preselected amount of bond area between
the two portions 22, 24 of the joint 26. Reduction of the length of the transition
section 25 may also reduce residual stresses in the joint 26 after the aircraft is
placed in service where large temperature extremes may be encountered during either
normal or extreme operations.
[0028] FIGS. 5 and 6 illustrate details of two adjoining layers 38 of the joint 26 shown
in FIGS. 2-4. In this example, each layer 38 comprises four plies 35 having a collective
total thickness
T1. The individual metal sheets 37 of the adjacent layers 38 are bonded together by
means of a layer of structural adhesive 45, which may comprise a commercial film adhesive
or other forms of a suitable adhesive that is placed between the metal sheets 36 during
the lay-up process.
[0029] The combined thickness of each metal sheet 37 and one layer of adhesive 45 represented
as
T2 in FIG. 5 is substantially equal to the thickness
T1 of the composite plies 35 in the layer 38. Although not shown in the Figures, a thin
film of adhesive may be placed between the plies 35 to increase the interlaminar bond
strength. In one practical embodiment, titanium alloy metal sheets 37 may be used
which each have a thickness of approximately 0.0025 inches, the film adhesive 45 may
be approximately 0.005 inches thick, and four composite carbon fiber epoxy plies 35
may be used in each layer 38 having a collective total thickness of about 0.30 inches.
Depending on the application, the use of metals other than titanium may be possible.
The distance between adjacent transition points 39, and thus the length of the overlap
between the layers 38, as well as the thickness and number of composite plies 35 and
the thickness of the metal sheets 37 will depend on the requirements of the particular
application, including the type and magnitude of the loads that are to be transmitted
through the joint 26, and possibly other performance specifications. It should be
noted here that the bonded metal sheets 37 is not limited to use in a composite metal
joint 26 discussed above. As will be discussed later below, a metal structure comprising
bonded metal sheets 37 has a variety of other applications because of the superior
strain performance it may exhibit, compared to monolithic metal structures.
[0030] The differing layers 38 of the joint 26 between the two differing materials of the
composite and metal portions 22, 24 respectively (FIG. 1), render the structure 20
well suited to nondestructive evaluations of bond quality using embedded or mounted
sensors (not shown). Ultrasonic structural waves (not shown) may be introduced into
the structure 20 at the edge of the metal portion 24, at the composite portion 22
or in the transition section 25. These ultrasonic waves travel through what amounts
to a waveguide formed by the metal 37 sheets and the interfaces (not shown) between
the composite plies 35 and the metal sheets 37. MEMS-based (microelectromechanical)
sensors, thin piezo-electric sensors (not shown) or other transducers placed in the
structure 20 may be used to receive the ultrasonic structural waves for purposes on
analyzing the condition of the bondlines in the joint 26.
[0031] Referring now to FIG. 7, one method of making the composite structure 20 comprises
forming a multi-layer composite lay-up as shown at 65. Forming the lay-up includes
laying up a composite resin portion 22 at step 67, and laying up a metal portion 24
at 69. The step 65 of forming the layup further includes forming a composite-to-metal
joint between the composite resin portion and the metal portion of the lay-up, shown
at 71.
[0032] FIG. 8 illustrates additional details of the method shown in FIG. 7. Beginning at
step 40, individual metal sheets 37 are trimmed to a desired size and/or shape. Next
at 42, the surfaces of the metal sheets 37 are prepared by suitable processes that
may include cleaning the sheets 37 with a solvent, drying them, etc. Then at 44, the
lay-up is assembled by laying up the metal sheets 36 and the composite plies 35 in
a sequence that is determined by a predefined ply schedule (not shown) which includes
a predetermined staggering of the transition points 39 between the plies 35 and the
metal sheet 37 in each layer 38.
[0033] During the lay-up process, the metal sheets 37 are sequenced like plies into the
lay-up, much like composite plies are sequenced into a lay-up in a conventional lay-up
process. As shown at step 46, adhesive may be introduced between the metal sheets
37 in order to bond them together into a unitized metal structure. Similarly, although
not shown in FIG. 8, a bonding adhesive may be introduced between the individual composite
plies 35 in order to increase the bond strength between these plies 35. Next, at 48,
the lay-up may be compacted using any of several known compaction techniques, such
as vacuum bagging following which the lay-up is cured at step 50 using autoclave or
out-of-autoclave curing processes. At step 52, the cured composite structure 20 may
be trimmed and/or inspected, as necessary.
[0034] FIG. 9 illustrates still another embodiment of a method of making a hybrid composite
part 20. The method begins at step 73 with laying at least one composite ply 35 that
is terminated at an interface location 39 on a suitable layup tool (not shown). At
75, an adjacent metal ply 37 is laid up which is substantially the same thickness
as the adjacent composite material ply 35. As shown at 77, the layup process is repeated
with a composite-to-metal interface 39 that is staggered toward the metal edge 24a
of the part 20 from the transition point 39. A 79, a ply 45 of structural adhesive
is laid between the metal plies 37. Steps 73-79 are repeated successively to produce
a nested splice 27 and a staggered interface stacking forming nested tabs 29 to the
full thickness of the hybrid part 20, with none of the composite plies 35 extending
fully to the metal edge 24a of the part 20. Although not shown in FIG. 9, the completed
layup is vacuum bagged processed to remove voids, and is subsequently cured using
any suitable curing method.
[0035] The composite-to-metal joint 26 previously described may be constructed in any of
a variety of joint configurations in which the composite material plies 35 are interleafed
with the metal plies 37. For example, referring to FIG. 10, the transition section
25 of the hybrid composite structure 20 may include a composite-to-metal joint 26
having a relatively shallow taper resulting from lengths
L of overlap between the composite and metal plies 35, 37 that are relatively long.
In the example shown in FIG. 10, the composite-to-metal joint 26 is a double tapered
finger joint. In comparison, as shown in FIG. 11, shorter lengths L of the overlap
between the composite and metal plies 35, 37 results in a double tapered finger joint
26 that has a relatively steep taper, in turn resulting in a shorter transition section
25 between the composite resin and metal portions 22, 24 respectively. The length
L of the overlap may be optimized for the particular application.
[0036] FIGS. 12-17 illustrate other examples of composite-to-metal joint 26 configurations.
In one alternative, the composite-to-metal joint 26 may comprise a double tapered
finger joint 26 that includes a tapered or layered multi-ply construction above and
below a composite-to-metal interface 39, wherein one or more overlap lengths, e.g.,
lengths L, may be chosen or optimized relative to a particular real estate constraint,
area, or transitional stress or strain requirement. In one example, the real estate
constraint or area may require a shorter transition section, for instance, between
the composite resin and metal portions. In some applications, a transitional stress
or strain requirement may require progressively less stress or strain along a portion
of the structure. For example, FIG. 12 illustrates a single taper lap joint 26, while
FIG. 13 illustrates a single reverse taper lap joint 26. In FIG. 14, the joint 26
is configured as a double tapered, substantially symmetrical, staggered finger lap
joint while FIG. 15 illustrates a reverse double tapered finger lap joint 26. The
use of the staggered finger lap joints 26 shown in FIGS. 14 and 15 may be preferred
in some applications because the joint may have a CTE interface that is less than
an equivalent step lap joint of a longer transition section 25 (FIG. 10). In FIG.
16, the composite-to-metal joint 26 takes the form of a vertical lap finger joint,
while FIG. 17 illustrates a composite-to-metal joint 26 in which the overlap between
the composite and the metal plies 35, 37 is variable through the thickness of the
joint 26.
[0037] Attention is now directed to FIGS. 18 and 19 which illustrate a hybrid composite
structure 20 comprising a composite resin portion 22 and a metal portion 24 that forms
a metal laminate reinforcement 76 around a fastener 78 passing through the hybrid
composite structure 20. The metal portion 24 forming the metal laminate reinforcement
76 comprises a stack 36 of metal sheets or plies 37 that are bonded together, similar
to the metal laminates previously described. The metal laminate reinforcement 76 is
connected to the surrounding composite resin portion 22 by a circumferential composite-to-metal
joint 26, as shown in FIG. 19 which, in the illustrated embodiment, comprises a double
tapered finger lap joint, similar to that shown in FIGS. 4, 10, 11 and 14. In one
alternative, staggered finger lap joints may include a transition region where one
or more edges of composite material plies, metal plies, or combinations thereof may
have varying levels of overlap or non-overlap to achieve or meet a desired CTE interface
coefficient, a desired real estate constraint, an area constraint, or transitional
stress or strain requirement. In one example, real estate constraint or area may require
a shorter transition section, for instance, between the composite resin and metal
portions or metal plies. In one example, transitional stress or strain requirement
may require progressively less stress or strain along a portion of the structure.
[0038] The metal laminate reinforcement 76 includes a central through-hole 85 through which
the fastener 78 passes. The fastener 78 may comprise for example and without limitation,
a bolt or rivet 78 having a body 78a and heads 78b and 78c. Although not shown in
the drawings, the fastener 78 may be used to attach a structure to the composite structure
20, or to secure the hybrid composite structure 20 to another structure. The metal
laminate reinforcement 76 functions to strengthen the area surrounding the fastener
78 and may better enable the composite structure 20 to carry loads in the area of
the fastener 78.
[0039] The composite-to-metal joint 26 previously described may be employed in a variety
of applications, including those in the aerospace industry to join composite structures,
especially in areas where a composite structure is highly loaded. For example, referring
to FIG. 20, an airplane 80 broadly comprises a fuselage 82, left and right wings 84,
a vertical stabilizer 92 and a pair of horizontal stabilizers 94, and a wing box 108.
The airplane 80 may further include a pair of engines 88 surrounded by engine nacelles
86, and landing gear 90.
[0040] The composite-to-metal joint 26 previously described may be employed to join or mount
any of the components shown in FIG. 20. For example, composite-to-metal joints 26
may be employed to mount the wings 84 on the center wing box 108, as will be discussed
below in more detail. Similarly, a composite-to-metal joint 26 may be employed to
attach the vertical stabilizer 92 and/or the horizontal stabilizers 94 to the fuselage
82. The composite-to-metal joints 26 may be employed to mount the landing gear 90
on the wings 84, as well as to mount engines 88 and engine nacelles 86 on pylons (not
shown) on the wings 84. Further, the disclosed composite-to-metal joint 26 may be
employed to join fuselage sections 82a together. For example, referring to FIGS. 20
and 21, fuselage sections 82a may be joined together by a co-bonded lap joint indicated
at 96, wherein each of the adjoining fuselage sections 82a comprises a metal laminate
stack 36 and finger overlaps 98, 100 between composite resin and metal plies 35, 37
respectively. In this example, the metal laminate stacks 36 of the respective fuselage
sections 82a may be joined together, as by bonding using a suitable bonding adhesive.
[0041] Referring now to FIG. 22, each of the wings 84 (FIG. 20) may be attached to the center
wing box 108 by an attachment joint, generally indicated at 104. Each of the wing
106 and the wing box 108 broadly comprises an outer skin 120 attached to spanwise
extending spars 110. The attachment joint 104 includes an attachment fitting 114 having
a pair of flanges 118 that are attached by bolts 122 or other suitable fasteners to
the skins 120. The attachment joint 104 may be reinforced by C-shaped channels 112
and brackets 116.
[0042] Referring also now to FIGS. 23 and 24, each of the skins 120 includes a metal portion
24 that also forms an integrated attachment fitting which is connected to a composite
resin portion 22 by a composite-to-metal joint 26 of the type previously described.
Although not shown in FIGS. 23 and 24, the metal portion 24 of the joint 26 is formed
by laminated metal plies 37, and the composite resin portion of the joint 26 is formed
by laminated composite resin plies 35. As particularly shown in FIG. 24, the metal
portion 24 of the joint 26 may be scarfed at 128 to receive one of the flanges 118
therein. Metal portions 24 include through-holes 124 that are aligned with the through-holes
126 in the flanges 118 of the fitting 116. It may thus be appreciated that attachment
joint 104 is reinforced by the presence of the metal portions 24 which are attached
to the metal attachment fitting 114 by the bolts 122.
[0043] FIGS. 25 and 26 illustrate another application of composite-to-metal joint 26 that
may be employed to attach a vertical stabilizer 92 or similar airfoil to an aircraft
fuselage 82. As shown in FIG. 25, the vertical stabilizer 92 may comprise a series
of generally upwardly extending spars 130 connected with ribs 132. A series of attachment
lugs 134 on the bottom of the stabilizer 92 are each attached to mounting ears 138
on the fuselage 82 by means of attachment bolts 136 received within bushings 140 in
the lugs 134. Each of the lugs 134 comprises a fiber reinforced composite resin portion
22 and a metal portion 24 which may comprise a metal laminate. The composite resin
portion 22 is joined to the metal portion 24 by a composite-to-metal joint 26 of the
type previously described. It may thus be appreciated that while the lug 134 is lightweight
because of its predominantly composite construction, the area at which the lug 134
is attached to the fuselage 82 comprises a metal portion 24 which has a load bearing
capacity that may be greater than the composite resin portion 22.
[0044] Attention is now directed to FIGS. 27 and 28 which illustrate the use of a composite-to-metal
joint 26 employed to reinforce the edges 142 of a fiber reinforced composite resin
structure, which in the illustrated example comprises the skin 120 of an aircraft
80. In this example, a fuselage hatch 141 has a periphery 142 terminating in an edge
144 (FIG. 28) that is reinforced by a metal portion 24 comprising a metal laminate
stack 36. The metal portion 24 is joined to the composite skin 120 by a composite-to-metal
joint 26, of the type previously described. In this example, the edge 24a of the metal
portion 24 defines the fuselage hatch 141 opening. The composite-to-fiber joint 26
may also be used to reinforce the skin 120 around other openings, such as cockpit
windows 125 and passenger windows 127.
[0045] Referring now to FIG. 29, the composite-to-metal joint 26 may be employed to attach
components on other types of aircraft, such as, for example and without limitation,
a helicopter 146. The helicopter 146 includes a main rotor assembly 148 and a tail
rotor assembly 150. The main rotor assembly 148 includes a plurality of main rotator
blades 152, and the tail rotor assembly 150 comprises a plurality of tail rotor blades
154. Each of the main rotor blades 152 is mounted on a rotor hub 156 secured to a
rotating mast 168 that is powered by one or more engines 160. Referring particularly
to FIGS. 30 and 31, each of the main rotor blades 152 is attached to the hub 156 by
means of blade grips 164. The root 162 of each blade 152 is held on the blade grips
164 by retention bolts 166. Each of the blades 152 includes an elongate outer composite
resin portion 22 which may be a carbon fiber epoxy composite, and a metal portion
24 that is attached to the blade grips 164 by the retention bolts 166. Metal portion
24 of the blade 152 is connected to the outer composite resin portion by a composite-to-metal
joint 26 of the type previously described. The tail rotor blades 154 shown in FIG.
29 may similarly be attached to the tail rotor assembly 150 by a composite-to-metal
joint 26.
[0046] Referring to FIG. 32, a metal laminate 170 comprises a plurality of generally flexible
metal sheets or plies 37 which are bonded together by layers 45 of a suitable adhesive
to form a structure that may exhibit performance properties that are superior to a
comparable monolithic metal structure. The layers 45 of adhesive may comprise a conventional
film-type structural adhesive. The metal plies 37 may be formed of the same metal
or may be formed of differing metals, depending on the particular application. When
the metal laminate 170 is placed in tension 175, the tension load is individually
directed to each of the metal laminate plies 37, thereby distributing the tension
load generally evenly throughout the metal structure 170. Thus, in the event of an
irregularity or inconsistency in one of the metal plies 37 that may reduce the load
carrying ability of the ply 37, the reduction is limited to that particular ply and
the applied tension load is redistributed to the remaining metal plies 37 which provide
strain relief. In other words, sensitive areas (i.e. plies 37) of the metal laminate
170 that are under load locally strain and transfer the load to adjacent metal plies
37, resulting in a form of a progressive loading of the metal laminate 170.
[0047] The metal laminate 170 shown in FIG. 32 may be employed to form composite-to-metal
joints 26 of the type previously described, but may have other applications as well.
For example, referring to FIG. 33, two generally flat metal laminates 170a, 170b may
be attached to each other by a lap joint 172 and fasteners 178 that pass through through-holes
173 the metal laminates 170a, 170b. The lap joint 172 employing may exhibit characteristics
that are superior to joints employing monolithic structures. The metal laminates 170a,
170b may form the edges of a composite structure to which the metal laminates 170a,
170b are joined by composite-to-metal joints 26 of the type previously described.
[0048] Referring to FIG. 34, a method of fabricating a structure begins at 180, with assembling
at least a first stack 36 of metal sheets or plies 37. The metal sheets or plies 37
are then laminated together at 182 by placing a layer of structural adhesive between
the sheets or plies 37 which bonds and laminates the sheets or plies 37 together into
a first metal laminate 170a. Then, optionally at 184, a second stack of metal sheets
or plies 37 is assembled and laminated together at 186 into a second metal laminate
170b. At 188, one or more through-holes 173 are formed in the first and second laminates
170a, 170b. At 190, fasteners are installed in the though-holes 173 to fasten the
metal laminates 170a, 170b together.
[0049] Referring to FIG. 35, selected areas of a fiber reinforced composite resin laminate
structure may be reinforced by a method that begins at step 192 with assembling a
metal laminate reinforcement 76. At step 194, composite resin plies 35 of the composite
resin laminate structure are interleafed with the metal laminate plies 37 of the metal
laminate reinforcement 76 to form a composite-to-metal step lap joint 26 in the area
of the composite resin laminate structure to be reinforced. As previously discussed,
the metal laminate reinforcement 76 may be used to reinforce an edge of the composite
resin laminate structure, or to provide a metal reinforced area around a fastener
78. Thus, optionally, at step 196, a through-hole 85 may be formed in the metal reinforcement
76, and at 198, a fastener 78 may be installed in the through-hole 85.
[0050] Embodiments of the disclosure may find use in a variety of potential applications,
particularly in the transportation industry, including for example, aerospace, marine
and automotive applications. Thus, referring now to FIGS. 36 and 37, embodiments of
the disclosure may be used in the context of an air10raft manufacturing and service
method 200 as shown in FIG. 36 and an aircraft 202 as shown in FIG. 37. Aircraft applications
of the disclosed embodiments may include, for example, a wide variety of structural
composite parts and components, especially those requiring local reinforcement and/or
the use of fasteners during the assembly process. During pre-production, exemplary
method 200 may include specification and design 204 of the aircraft 202 and material
procurement 206. During production, component and subassembly manufacturing 208 and
system integration 210 of the aircraft 202 takes place. Thereafter, the aircraft 202
may go through certification and delivery 212 in order to be placed in service 214.
While in service by a customer, the aircraft 202 is scheduled for routine maintenance
and service 216.
[0051] Each of the processes of method 200 may be performed or carried out by a system integrator,
a third party, and/or an operator (e.g., a customer). For the purposes of this description,
a system integrator may include without limitation any number of aircraft manufacturers
and major-system subcontractors; a third party may include without limitation any
number of vendors, subcontractors, and suppliers; and an operator may be an airline,
leasing company, military entity, service organization, and so on.
[0052] As shown in FIG. 37, the aircraft 202 produced by exemplary method 200 may include
an airframe 218 with a plurality of systems 220 and an interior 222. Examples of high-level
systems 220 include one or more of a propulsion system 224, an electrical system 226,
a hydraulic system 228, and an environmental system 230. Any number of other systems
may be included. The disclosed method may be employed to fabricate parts, structures
and components used in the airframe 218 or in the interior 222. Although an aerospace
example is shown, the principles of the disclosure may be applied to other industries,
such as the marine and automotive industries.
[0053] Systems and methods embodied herein may be employed during any one or more of the
stages of the production and service method 200. For example, parts, structures and
components corresponding to production process 208 may be fabricated or manufactured
in a manner similar to parts, structures and components produced while the aircraft
200 is in service. Also, one or more apparatus embodiments, method embodiments, or
a combination thereof may be utilized during the production stages 208 and 210, for
example, by substantially expediting assembly of or reducing the cost of an aircraft
200. Similarly, one or more of apparatus embodiments, method embodiments, or a combination
thereof may be utilized while the aircraft 202 is in service, for example and without
limitation, to maintenance and service 216.
[0054] Although the embodiments of this disclosure have been described with respect to certain
exemplary embodiments, it is to be understood that the specific embodiments are for
purposes of illustration and not limitation, as other variations will occur to those
of skill in the art.